EP1801361A1 - Ventilator- oder Verdichtergehäuse - Google Patents

Ventilator- oder Verdichtergehäuse Download PDF

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Publication number
EP1801361A1
EP1801361A1 EP06255950A EP06255950A EP1801361A1 EP 1801361 A1 EP1801361 A1 EP 1801361A1 EP 06255950 A EP06255950 A EP 06255950A EP 06255950 A EP06255950 A EP 06255950A EP 1801361 A1 EP1801361 A1 EP 1801361A1
Authority
EP
European Patent Office
Prior art keywords
casing
fan
grooves
helical groove
blades
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP06255950A
Other languages
English (en)
French (fr)
Inventor
Larry Collins
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce PLC
Original Assignee
Rolls Royce PLC
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Rolls Royce PLC filed Critical Rolls Royce PLC
Publication of EP1801361A1 publication Critical patent/EP1801361A1/de
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/522—Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps
    • F04D29/526—Details of the casing section radially opposing blade tips
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F04D29/685—Inducing localised fluid recirculation in the stator-rotor interface
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/14—Casings or housings protecting or supporting assemblies within
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/25—Three-dimensional helical

Definitions

  • the invention relates to a fan or compressor casing and more particularly to a modified fan or compressor casing for having improved surge margin performance.
  • the invention may be utilised in a fan section at the front of a gas turbine engine, or in an axial flow compressor section of the engine, or in an independent rotary fan arrangement for example of the kind used as a lift fan for a vertical / short take off aircraft.
  • Such arrangements have in common a rotary stage enclosed within a generally cylindrical casing.
  • the aerodynamic performance of such arrangements and the phenomenon of surge are well understood. It is known to increase their surge margin by adopting certain modifications to the casing wall in the vicinity of the rotary stage that allow limited re-circulation of air over the tips of the blades.
  • One such modification for improving the surge margin of a rotary stage of this kind uses a number of circumferential grooves formed in the casing wall. In operation, as the stage rotates a certain amount of air passing through the stage is spilled from the blade tips into the grooves and can travel around the groove according to pressure distribution therein. The drawback in such an arrangement is that foreign objects are also centrifuged into the grooves and can accumulate therein leading to impaired performance of the modification. This is undesirable as it leads inevitably to a reduction of the available surge margin in the rotary stage.
  • the present invention solves the above-mentioned drawback by avoiding the accumulation of foreign objects in the aforementioned tip treatment grooves.
  • a gas turbine engine driven fan or compressor stage casing comprising a generally cylindrical casing portion encircling a fan or compressor rotor having a multiplicity of blades spaced apart around the rotor, the casing portion having formed in its inner surface over the tips of the blades at least one helical groove characterised in that the at least one helical groove has a gradually reducing depth in the downstream direction.
  • the at least one helical groove extends for a plurality of turns around the inner surface of the casing, and in addition to the depth of groove or grooves being tapered towards the downstream direction, the grooved portion may also extend beyond the trailing edges of the rotor blades.
  • the groove or grooves may also be tapered towards the forward end and the grooved portion may extend ahead of the leading edges of the rotor blades.
  • FIG. 1 there is shown a generally cylindrical portion 2 of an axial flow compressor fan casing, in this case the fan casing forms part of lift fan arrangement for a V/STOL aircraft but it could equally form part of the intake to a gas turbine propulsion engine.
  • the casing 2 incorporates integral annular flanges 4,6 by means of which the casing may be attached to adjacent lift fan or engine sections in axial flow series as appropriate.
  • a tip portion of a blade forming part of an associated rotary fan is shown at 8 to illustrate the close spacing between the two parts.
  • the radially outer tip surfaces 10 of the blades 8 are positioned close to the inner surface 12 of the casing 2.
  • the exact size of the tip clearance gap varies with rotor speed mainly due to elongation of the blade radius as a result mainly of the effect of centrifugal forces and blade untwist.
  • the compressor, or fan, working line may be forced beyond the stability line leading to a surge.
  • the stage is designed such that in normal operating range a surge margin is maintained between the working and stability lines sufficient to avoid surge.
  • the situations that produce surge and the transient operating conditions during a surge event are well known, and it is not intended to analyse the several possible types of surge in depth here.
  • the inlet mass flow of the rotary stage stalls or varies with time as the flow rate oscillates and may even become negative.
  • Certain design steps may be taken to avoid or alleviate the conditions that give rise to a stall.
  • One such step involves the design of the casing wall in the vicinity of the blade tips.
  • the grooves 14a-j are cut into a demountable, liner section 16 secured in a cut-out formed in the inner surface 12 of the casing section 2.
  • the liner section 16 may be made from different material compared to the remainder of the casing 2. It may be made in a number of segments and secured to the casing by appropriate means.
  • FIG. 2 the circular grooves 14a-j of the prior art ( Figure 1) are replaced by at least one or a single helical groove 18 cut into the liner section 16.
  • the helix or helical groove 18 is handed to suit the direction of rotation of the rotor and blades 8. Air entrained in the groove and foreign particles are forced to travel along the helix from their point of entry and are expelled at the end of the groove.
  • the direction of the helix may be chosen so that particles are transported either upstream or downstream relative to their entry point into the groove.
  • helical grooves 18 There may be a plurality of such helical grooves 18 arranged to commence at a like number of different points 20 spaced apart around the circumference of the casing.
  • particles or foreign objects will tend to be expelled from the groove in one circumferential position, basically at the downstream end 22 of the groove but factors such as orientation may be an influence. It is considered undesirable to concentrate potentially abrasive particles in one position as it could result in increased localised abrasion of components downstream of the ejection point against which the debris is incident. Therefore it is preferred to employ a plurality of helical grooves 18 in the liner 16.
  • Each of these helical grooves has the same number of turns and is of the same length, and the start points 20 or upstream ends are equidistantly spaced apart around the circumference of the liner 16.
  • the exit or downstream points 22 are correspondingly spaced apart around the circumference of the liner surface 16.
  • the termination of each groove 18 is such that foreign particles are deflected away from adjacent, downstream components.
  • the depth of the helical grooves 18 is progressively reduced towards the chosen exit point 22 at least to reduce the chance that particles may become trapped at an abrupt end as in the prior art. It is envisaged that normally, that is in most cases only the depth of a portion of a helical grove 18 towards the downstream end 22 would be formed with a progressively reducing depth, as illustrated in Figure 2. However, the upstream ends 20 of the or each helix or helical groove 18 may be tapered in similar fashion, that is the helical groove 18 may be formed with a progressively decreasing depth in an upstream direction, as illustrated in Figure 3.
  • Multiple helical grooves 18 are more efficient at expelling particles, as the pitch of each helix 18 increases in proportion to the number of grooves, and less abrasion takes place. With less abrasion, or erosion, an optimum design geometry is maintained for a longer period so that the level of aerodynamic performance is preserved for a longer period of time compared to the existing prior art arrangement.
  • the grooves 18 serve a dual purpose. In addition to the grooves 18 functioning to collect and transport through the stage ingested foreign object particles, their primary function is to provide a re-circulation path in the vicinity of the tips of the rotor blades 8 to increase the stability and surge margin of the rotor stage in known fashion. Therefore, as is well known in that particular art the start 20 of the re-circulation grooves 18 may extend upstream of the leading edges 24 of the blades 8, and the exit 22 of the grooves 18 may extend downstream of the trailing edges 26 of the blades 8, as illustrated in Figure 4 of the drawings.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP06255950A 2005-12-22 2006-11-21 Ventilator- oder Verdichtergehäuse Withdrawn EP1801361A1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GBGB0526011.2A GB0526011D0 (en) 2005-12-22 2005-12-22 Fan or compressor casing

Publications (1)

Publication Number Publication Date
EP1801361A1 true EP1801361A1 (de) 2007-06-27

Family

ID=35840873

Family Applications (1)

Application Number Title Priority Date Filing Date
EP06255950A Withdrawn EP1801361A1 (de) 2005-12-22 2006-11-21 Ventilator- oder Verdichtergehäuse

Country Status (3)

Country Link
US (1) US20070147989A1 (de)
EP (1) EP1801361A1 (de)
GB (1) GB0526011D0 (de)

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2929349A1 (fr) * 2008-03-28 2009-10-02 Snecma Sa Carter pour roue a aubes mobiles de turbomachine
EP2484913A2 (de) 2011-02-03 2012-08-08 Rolls-Royce plc Turbomaschine umfassend ein Ringgehäuse und einen Schaufelrotor
EP2687684A1 (de) * 2012-07-17 2014-01-22 MTU Aero Engines GmbH Anstreifbelag mit Spiralnuten in einer Strömungsmaschine
FR2994718A1 (fr) * 2012-08-27 2014-02-28 Snecma Carter a traitements de carter arasants
EP2604805A3 (de) * 2011-12-15 2015-08-19 General Electric Company Statoraussenringanordnung für ein Gasturbinentriebwerk
US9874104B2 (en) 2015-02-27 2018-01-23 General Electric Company Method and system for a ceramic matrix composite shroud hanger assembly
US10309244B2 (en) 2013-12-12 2019-06-04 General Electric Company CMC shroud support system
US10378387B2 (en) 2013-05-17 2019-08-13 General Electric Company CMC shroud support system of a gas turbine
US10400619B2 (en) 2014-06-12 2019-09-03 General Electric Company Shroud hanger assembly
US10465558B2 (en) 2014-06-12 2019-11-05 General Electric Company Multi-piece shroud hanger assembly
EP4006313A1 (de) * 2018-12-28 2022-06-01 Honeywell International Inc. Verdichterabschnitt eines gasturbinentriebwerks mit hybridummantelung mit gehäusebehandlung und abreibbarem abschnitt
US11668207B2 (en) 2014-06-12 2023-06-06 General Electric Company Shroud hanger assembly

Families Citing this family (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7679906B2 (en) * 2007-11-05 2010-03-16 Microsoft Corporation Liquid resistant A/C adaptor
US8177494B2 (en) * 2009-03-15 2012-05-15 United Technologies Corporation Buried casing treatment strip for a gas turbine engine
US8602720B2 (en) 2010-06-22 2013-12-10 Honeywell International Inc. Compressors with casing treatments in gas turbine engines
FR2988146B1 (fr) * 2012-03-15 2014-04-11 Snecma Carter pour roue a aubes de turbomachine ameliore et turbomachine equipee dudit carter
FR2995949B1 (fr) * 2012-09-25 2018-05-25 Safran Aircraft Engines Carter de turbomachine
WO2014158236A1 (en) * 2013-03-12 2014-10-02 United Technologies Corporation Cantilever stator with vortex initiation feature
TW201518607A (zh) * 2013-11-14 2015-05-16 Hon Hai Prec Ind Co Ltd 風扇
US10465716B2 (en) 2014-08-08 2019-11-05 Pratt & Whitney Canada Corp. Compressor casing
US10066640B2 (en) * 2015-02-10 2018-09-04 United Technologies Corporation Optimized circumferential groove casing treatment for axial compressors
US10107307B2 (en) 2015-04-14 2018-10-23 Pratt & Whitney Canada Corp. Gas turbine engine rotor casing treatment
WO2016168528A1 (en) 2015-04-15 2016-10-20 Robert Bosch Gmbh Free-tipped axial fan assembly
US9951642B2 (en) * 2015-05-08 2018-04-24 United Technologies Corporation Intermittent grooved soft abradable material to reduce blade tip temperature
GB201515934D0 (en) * 2015-09-09 2015-10-21 Rolls Royce Plc Method of manufacturing a gas turbine engine
US10428674B2 (en) * 2017-01-31 2019-10-01 Rolls-Royce North American Technologies Inc. Gas turbine engine features for tip clearance inspection
US10648484B2 (en) * 2017-02-14 2020-05-12 Honeywell International Inc. Grooved shroud casing treatment for high pressure compressor in a turbine engine
US11242769B2 (en) * 2018-12-17 2022-02-08 Raytheon Technologies Corporation Additively controlled surface roughness for designed performance
US11136895B2 (en) * 2019-02-06 2021-10-05 Raytheon Technologies Corporation Spiraling grooves as a hub treatment for cantilevered stators in compressors
US11015465B2 (en) 2019-03-25 2021-05-25 Honeywell International Inc. Compressor section of gas turbine engine including shroud with serrated casing treatment
EP3885537A1 (de) * 2020-03-26 2021-09-29 Unison Industries LLC Haltesystem für luftturbinenstarter und verfahren zur herstellung solches systems
US20230151825A1 (en) * 2021-11-17 2023-05-18 Pratt & Whitney Canada Corp. Compressor shroud with swept grooves
US11970985B1 (en) 2023-08-16 2024-04-30 Rolls-Royce North American Technologies Inc. Adjustable air flow plenum with pivoting vanes for a fan of a gas turbine engine
US11965528B1 (en) 2023-08-16 2024-04-23 Rolls-Royce North American Technologies Inc. Adjustable air flow plenum with circumferential movable closure for a fan of a gas turbine engine

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0151071A2 (de) * 1984-02-01 1985-08-07 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." Dichtungssystem zwischen den Schaufelspitzen und dem Gehäuse eines Axialkompressors
GB2245312A (en) * 1984-06-19 1992-01-02 Rolls Royce Plc Axial flow compressor surge margin improvement
EP1101947A2 (de) * 1999-11-15 2001-05-23 General Electric Company Abriebbeständige Verdichterstufe
EP1243797A2 (de) * 2001-03-19 2002-09-25 Williams International Co., L.L.C. Verdichtergehäuse für eine Gasturbine

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US3261228A (en) * 1964-04-02 1966-07-19 United Aircraft Corp Disk fragment energy absorption and containment means
GB1518293A (en) * 1975-09-25 1978-07-19 Rolls Royce Axial flow compressors particularly for gas turbine engines
FR2406074A1 (fr) * 1977-10-11 1979-05-11 Snecma Dispositif de securite pour machine tournante axiale
US6302643B1 (en) * 1999-04-26 2001-10-16 Hitachi, Ltd. Turbo machines
US6290458B1 (en) * 1999-09-20 2001-09-18 Hitachi, Ltd. Turbo machines
GB0216952D0 (en) * 2002-07-20 2002-08-28 Rolls Royce Plc Gas turbine engine casing and rotor blade arrangement

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0151071A2 (de) * 1984-02-01 1985-08-07 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." Dichtungssystem zwischen den Schaufelspitzen und dem Gehäuse eines Axialkompressors
GB2245312A (en) * 1984-06-19 1992-01-02 Rolls Royce Plc Axial flow compressor surge margin improvement
EP1101947A2 (de) * 1999-11-15 2001-05-23 General Electric Company Abriebbeständige Verdichterstufe
EP1243797A2 (de) * 2001-03-19 2002-09-25 Williams International Co., L.L.C. Verdichtergehäuse für eine Gasturbine

Cited By (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009125135A3 (fr) * 2008-03-28 2009-12-10 Snecma Carter pour roue a aubes mobiles de turbomachine
RU2491447C2 (ru) * 2008-03-28 2013-08-27 Снекма Кожух для рабочего колеса турбомашины
FR2929349A1 (fr) * 2008-03-28 2009-10-02 Snecma Sa Carter pour roue a aubes mobiles de turbomachine
US8777558B2 (en) 2008-03-28 2014-07-15 Snecma Casing for a moving-blade wheel of turbomachine
EP2484913A2 (de) 2011-02-03 2012-08-08 Rolls-Royce plc Turbomaschine umfassend ein Ringgehäuse und einen Schaufelrotor
EP2604805A3 (de) * 2011-12-15 2015-08-19 General Electric Company Statoraussenringanordnung für ein Gasturbinentriebwerk
US9726043B2 (en) 2011-12-15 2017-08-08 General Electric Company Mounting apparatus for low-ductility turbine shroud
EP2687684A1 (de) * 2012-07-17 2014-01-22 MTU Aero Engines GmbH Anstreifbelag mit Spiralnuten in einer Strömungsmaschine
FR2994718A1 (fr) * 2012-08-27 2014-02-28 Snecma Carter a traitements de carter arasants
US10378387B2 (en) 2013-05-17 2019-08-13 General Electric Company CMC shroud support system of a gas turbine
US10309244B2 (en) 2013-12-12 2019-06-04 General Electric Company CMC shroud support system
US10400619B2 (en) 2014-06-12 2019-09-03 General Electric Company Shroud hanger assembly
US10465558B2 (en) 2014-06-12 2019-11-05 General Electric Company Multi-piece shroud hanger assembly
US11092029B2 (en) 2014-06-12 2021-08-17 General Electric Company Shroud hanger assembly
US11668207B2 (en) 2014-06-12 2023-06-06 General Electric Company Shroud hanger assembly
US9874104B2 (en) 2015-02-27 2018-01-23 General Electric Company Method and system for a ceramic matrix composite shroud hanger assembly
EP4006313A1 (de) * 2018-12-28 2022-06-01 Honeywell International Inc. Verdichterabschnitt eines gasturbinentriebwerks mit hybridummantelung mit gehäusebehandlung und abreibbarem abschnitt
US11421544B2 (en) 2018-12-28 2022-08-23 Honeywell International Inc. Compressor section of gas turbine engine including hybrid shroud with casing treatment and abradable section

Also Published As

Publication number Publication date
US20070147989A1 (en) 2007-06-28
GB0526011D0 (en) 2006-02-01

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